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  • Other Sources  (42)
  • Fluid Mechanics and Heat Transfer  (36)
  • Getreide
  • 1960-1964  (37)
  • 1925-1929  (5)
  • 1961  (37)
  • 1925  (5)
  • 1
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    Eine Getreidelaufkäferplage (Zabru tenebrioides) (1925)
    In:  Die kranke Pflanze 2, p. 117-119
    Details
    Publication Date: 1925
    Description: Allgemeine Beobachtungen zum Auftreten und zur Wanderung des Getreidelaufkäfers und dessen Larven; Nennung von Möglichkeiten zur Bekämpfung des Insekts KATASTER-BESCHREIBUNG: Abhängigkeit der Stärke des Auftretens von der Temperatur KATASTER-DETAIL: Delta T+, dann Auftreten der Larven +
    Keywords: Sachsen ; Beginn 20. Jahrhundert ; Insekten ; Boden ; Ertrag ; Getreide ; Landwirtschaft ; Pflanzenschädling ; Roggen ; Temperatur ; Weizen
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  • 2
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    Beobachtungen über die diesjährige Dürre (1925)
    In:  Pflanzenbau, Halbmonatsschrift für Saatwesen, Anbau und Pflege der Kulturpflanzen. II. Jahrgang 1925/26 S.106-108.
    Details
    Publication Date: 1925
    Description: Einfluss der Trockenheit im Sommer auf Getreide, Leguminosen und Zuckerrüben KATASTER-BESCHREIBUNG: Einfluss der Temperatur und des Niederschlags auf den Ertrag KATASTER-DETAIL: Delta T (Vorjahreswinter) + und Delta Nied (Sommer) -, dann Pflanzenschädlinge (Erbsenwickler, Fritfliege) +; Delta T (Vorjahreswinter) + und Delta Nied (Sommer) -, dann Ertrag (Kartoffeln) +; Delta T (Vorjahreswinter) + und Delta Nied (Sommer) -, dann Ertrag (Sommergerste) -;
    Keywords: Pommern ; 1925 ; Ertrag ; Getreide ; Niederschlag ; Trockenheit
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  • 3
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    Die Abhängigkeit der Düngerwirkung von der Witterung im Dikopshofer Dauerdüngungsversuch 1906–1957 (1961)
    In:  Zeitschrift für Pflanzenernährung, Düngung, Bodenkunde 95:1-9
    Details
    Publication Date: 1961
    Description: Witterung, insb. Niederschlag als Einflußgröße für Erträge, Auswertung von N- und P-Mangelversuchen KATASTER-BESCHREIBUNG: Einfluss der Witterung (Niederschlag) auf die Erträge bei Nährstoffmangel (P, K, N) KATASTER-DETAIL: Delta Nied -, dann P-Mangel + und Kali-Mangel +; Delta Nied +, dann Erträge (Weizen und Hafer) -; Delta Nied (Winter) +/-: Nied (Durchschnitt)= 399 mm, dann Ertrag bei N-Mangel + Nied (Durchschnitt)= 267 mm, dann Ertrag bei N-Mangel + Nied (Durchschnitt)= 327 mm, dann Ertrag bei N-Mangel -
    Keywords: Kölner Bucht ; 1906–1957 ; Ertrag ; Getreide ; Niederschlag ; Temperatur ; Witterung ; Düngung
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  • 4
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    Der Einfluß des Klimas auf die Erntebeschaffenheit des Getreides (1925)
    In:  Mitt. Deut. Landw. Gesell. 40:950-55.
    Details
    Publication Date: 1925
    Description: Zusammenhang zwischen Weizenqualität und den klimatischen Verhältnissen, Schwerpunkt auf die Physiologie des Samens und Qualität (Backfähigkeit) KATASTER-BESCHREIBUNG: KATASTER-DETAIL:
    Keywords: Europa ; letzten100 Jahre ; Ertrag ; Getreide ; Klima ; Niederschlag ; Temperatur ; Weizen
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  • 5
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    Zwei verbreitete Weizenschädlinge dieses Jahres (1925)
    In:  Sächs. Landw. Zeitschr. 73, p.614
    Details
    Publication Date: 1925
    Description: Beschreibung der Schäden und Erscheinungsformen des Weizenhalmtöters und der Weizenhalmfliege im Jahre 1926 in Sachsen. Nur bedingte Verknüpfung mit klimatischen Parametern. KATASTER-BESCHREIBUNG: KATASTER-DETAIL:
    Keywords: Sachsen ; 1926 ; Getreide ; Pflanzenschädling ; Weizen
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  • 6
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    Insektenschädigungen am Getreide (1925)
    In:  Georgine, Jahrgang 102, Nr. 68, p. 806-807
    Details
    Publication Date: 1925
    Description: Biologie und Bekämpfung von Fritfliege, Getreideblumenfliege, scheckige oder gelbe Halmfliege, Weizenfliege, Hessenfliege KATASTER-BESCHREIBUNG: Einfluss der Witterung (Temperatur, Niederschlag und Wind) auf die Schädlinge KATASTER-DETAIL: Delta T (Winter) +, Anzahl Maden der Halmfliege +; Delta Nied (September) + und Delta Wind (September) +, dann Hessenfliege -
    Keywords: Deutschland ; 1903-1925 ; Insekten ; Anbautermine ; Boden ; Ertrag ; Getreide ; Hafer ; Niederschlag ; Pflanzenschädling ; Temperatur ; Vegetationsperiode ; Weizen ; Wind ; Witterung ; Düngung ; Gerste
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  • 7
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    Laminar Heat-Transfer and Pressure-Distribution Studies on a Series of Reentry Nose Shapes at a Mach Number of 19.4 in Helium (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An experimental investigation has been conducted in the 2-inch helium tunnel at the Langley Research Center at a Mach number of 19.4 to determine the pressure distributions and heat-transfer characteristics of a family of reentry nose shapes. The pressure and heat-transfer-rate distributions on the nose shapes are compared with theoretical predictions to ascertain the limitations and validity of the theories at hypersonic speeds. The experimental results were found to be adequately predicted by existing theories. Two of the nose shapes were tested with variable-length flow-separation spikes. The results obtained by previous investigators of spike-nose bodies were found to prevail at the higher Mach number of the present investigation.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-891 , L-1345
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  • 8
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    An Experimental Investigation of the Damping of Liquid Oscillations in an Oblate Spheroidal Tank With and Without Baffles (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An experimental investigation was conducted to determine the damping of the fundamental antisymmetric mode of oscillation of liquids contained in an oblate spheroidal tank. The decay of the fundamental mode was studied for a range of liquid depths in tanks with and without baffles. In the investigation of baffle effects, ring and cruciform baffles of various sizes were fixed at different locations within the tank. Data presented show the variation of the damping factor with tank fullness and with baffle type, width, location, and orientation as well as the effects of the amplitude of the liquid oscillations and of small variations in the liquid kinematic viscosity on the damping factor. The results of the investigation indicate that the addition of ring baffles to the tank results in an increase in the available effective damping when the baffle plane is in a region near the equilibrium liquid surface, and that cruciform baffles are effective in the damping of the fundamental mode in the near-empty tank. No apparent changes in damping for the tanks having ring baffles were observed as the kinematic viscosity of the liquid was varied over a small range.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-808 , L-1348
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  • 9
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    Measurement and Empirical Correlation of Transpiration-Cooling Parameters on a 25 degree Cone in a Turbulent Boundary Layer in Both Free Flight and a Hot-Gas Jet (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Transpiration-cooling parameters are presented for a turbulent boundary layer on a cone configuration with a total angle of 250 which was tested in both free flight and in an ethylene-heated high-temperature jet at a Mach number of 2.0. The flight-tested cone was flown to a maximum Mach number of 4.08 and the jet tests were conducted at stagnation temperatures ranging from 937 R to 1,850 R. In general, the experimental heat transfer was in good agreement with the theoretical values. Inclusion of the ratio of local stream temperature to wall temperature in the nondimensional flow rate parameter enabled good correlation of both sets of transpiration data. The measured pressure at the forward station coincided with the theoretical pressure over a sharp cone; however, the measured pressure increased with distance from the nose tip.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-967 , L-1711
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  • 10
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    Experimental Determination of the Recovery Factor and Analytical Solution of the Conical Flow Field for a 20 deg Included Angle Cone at Mach Numbers of 4.6 and 6.0 and Stagnation Temperatures to 2600 degree R (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The local recovery factor was determined experimentally along the surface of a thin-walled 20 deg included angle cone for Mach numbers near 6.0 at stagnation temperatures between 1200 deg R and 2600 deg R. In addition, a similar cone configuration was tested at Mach numbers near 4.5 at stagnation temperatures of approximately 612 deg R. The local Reynolds number based on flow properties at the edge of the boundary layer ranged between 0.1 x 10(exp 4) and 3.5 x 10(exp 4) for tests at temperatures above 1200 deg R and between 6 x 10(exp 4) and 25 x 10(exp 4) for tests at temperatures near 612 deg R. The results indicated, generally, that the recovery factor can be predicted satisfactorily using the square root of the Prandtl number. No conclusion could be made as to the necessity of evaluating the Prandtl number at a reference temperature given by an empirical equation, as opposed to evaluating the Prandtl number at the wall temperature or static temperature of the gas at the cone surface. For the tests at temperatures above 1200 deg R (indicated herein as the tests conducted in the slip-flow region), two definite trends in the recovery data were observed - one of increasing recovery factor with decreasing stagnation pressure, which was associated with slip-flow effects and one of decreasing recovery factor with increasing temperature. The true cause of the latter trend could not be ascertained, but it was shown that this trend was not appreciably altered by the sources of error of the magnitude considered herein. The real-gas equations of state were used to determine accurately the local stream properties at the outer edge of the boundary layer of the cone. Included in the report, therefore, is a general solution for the conical flow of a real gas using the Beattie-Bridgeman equation of state. The largest effect of temperature was seen to be in the terms which were dependent upon the internal energy of the gas. The pressure and hence the pressure drag terms were unaffected.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-353 , A-318
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  • 11
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    Effects of Mach Number, Leading-Edge Bluntness, and Sweep on Boundary-Layer Transition on a Flat Plate (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The effects of leading-edge bluntness and sweep on boundary-layer transition on flat plate models were investigated at Mach numbers of 2.00, 2.50, 3.00, and 4.00. The effect of sweep on transition was also determined on a flat plate model equipped with an elliptical nose at a Mach number of 0.27. Models used for the supersonic investigation had leading-edge radii varying from 0.0005 to 0.040 inch. The free-stream unit Reynolds number was held constant at 15 million per foot for the supersonic tests and the angle of attack was 0 deg. Surface flow conditions were determined by visual observation and recorded photographically. The sublimation technique was used to indicate transition, and the fluorescent-oil technique was used to indicate flow separation. Measured Mach number and sweep effects on transition are compared with those predicted from shock-loss considerations as described in NACA Rep. 1312. For the models with the blunter leading edges, the transition Reynolds number (based on free-stream flow conditions) was approximately doubled by an increase in Mach number from 2.50 to 4.00; and nearly the same result was predicted from shock-loss considerations. At all super- sonic Mach numbers, increases in sweep reduced the transition Reynolds number and the amount of reduction increased with increases in bluntness. The shock-loss method considerably underestimated- the sweep effects, possibly because of the existence of crossflow instability associated with swept wings. At a Mach number of 0.27, no reduction in the transition Reynolds number with sweep was measured (as would be expected with no shock loss) until the sweep angle was attained where crossflow instability appeared.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1071 , A-481
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  • 12
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    Effects of Surface Recombination on Heat Transfer to Bodies in a High Enthalpy Stream of Partially Dissociated Nitrogen (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Heat-transfer rates to two surfaces having widely different catalytic effectiveness are compared at a Mach number of 6 in a low-density stream of partially dissociated nitrogen. The heat-transfer rate to a polished copper cylinder is twice as great as the heat-transfer rate to a silicon-monoxide-coated cylinder when the stream total energy content is 9000 Btu/lb. Various methods for determining the stream energy content, the stream velocity, and the stream Mach number have been developed and compared. It is shown that methods for estimating the stream energy content by means of purely aerodynamic concepts may neglect the sizable fraction of the stream energy contained in molecular dissociation.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1146 , A-1470
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  • 13
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    On the Induced Flow of an Electrically Conducting Liquid in a Rectangular Duct by Electric and Magnetic Fields of Finite Extent (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Reported here are the results of a systematic study of a model of the direct-current electromagnetic pump. Of particular interest is the motion imparted to the electrically conducting fluid in the rectangular duct by the body forces that result from applied electric and magnetic fields. The purpose of the investigation is to associate the observed fluid motion with the characteristics of the electric and magnetic fields which cause them. The experiments were carried out with electromagnetic fields that moved a stream of copper sulphate solution through a clear plastic channel. Ink filaments injected into the stream ahead of the region where the fields were applied identify the motion of the fluid elements as they passed through the test channel. Several magnetic field configurations were employed with a two-dimensional electric current distribution in order to study and identify the magnitude of some of the effects on the fluid motion brought about by nonuniformities in the electromagnetic fields. A theoretical analysis was used to guide and evaluate the identification of the several fluid motions observed. The agreement of the experimental data with the theoretical predictions is satisfactory. It is found that sizable variations in the velocity profile and pressure head of the output stream are produced by the shape of the electric and magnetic fields.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-347 , A-276
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  • 14
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    Laminar Boundary Layer Behind a Strong Shock Moving into Air (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The laminar wall boundary layer behind a strong shock advancing into stationary air has been determined. Numerical results have been obtained for shock Mach numbers up to 14 using real gas values for density and viscosity and assuming Prandtl and Lewis numbers of 0.72 and 1, respectively. The numerical results for shear and heat transfer agree, within 4 percent, with a previously presented approximate analytical expression for these quantities. A slight modification of this expression results in agreement with the numerical data to within 2.5 percent. Analytical expressions for boundary-layer thickness and displacement thickness, correct to within 4 percent for the present data, have also been obtained.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-291 , E-975
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  • 15
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    Radiant Heat Transfer Between Nongray Parallel Plates of Tungsten (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Net radiant heat flow between two infinite, parallel, tungsten plates was computed by summing the monochromatic energy exchange; the results are graphically presented as a function of the temperatures of the two surfaces. In general these fluxes range from approximately a to 25 percent greater than the results of gray-body computations based on the same emissivity data. The selection of spectral emissivity data and the computational procedure are discussed. The present analytical procedure is so arranged that, as spectral emissivity data for a material become available, these data can be readily introduced into the NASA data-reduction equipment, which has been programmed to compute the net heat flux for the particular geometry and basic assumptions cited in the text. Nongray-body computational techniques for determining radiant heat flux appear practical provided the combination of select spectral emissivity data and the proper mechanized data-reduction equipment are brought to bear on the problem.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1088 , E-1277
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  • 16
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    Pressure Loads Produced on a Flat-Plate Wing By Rocket Jets Exhausting in a Spanwise Direction Below the Wing and Perpendicular to a Free-Stream Flow of Mach Number 2.0 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation at a Reynolds number per foot of 14.4 x 10(exp 6) was made to determine the pressure loads produced on a flat-plate wing by rocket jets exhausting in a spanwise direction beneath the wing and perpendicular to a free-stream flow of Mach number 2.0. The ranges of the variables involved were (1) nozzle types - one sonic (jet Mach number of 1.00), two supersonic (jet Mach numbers of 1.74 and 3.04),. and one two-dimensional supersonic (jet Mach number of 1.71); (2) vertical nozzle positions beneath the wing of 4, 8 and 12 nozzle-throat diameters; and (3) ratios of rocket-chamber total pressure to free- stream static pressure from 0 to 130. The incremental normal force due to jet interference on the wing varied from one to two times the rocket thrust and generally decreased as the pressure ratio increased. The chordwise coordinate of the incremental-normal-force center of pressure remained upstream of the nozzle center line for the nozzle positions and pressure ratios of the investigation. The chordwise coordinate approached zero as the jet vertical distance beneath the wing increased. In the spanwise direction there was little change due to varying rocket-jet position and pressure ratio. Some boundary-layer flow separation on the wing was observed for the rocket jets close to the wing and at the higher pressure ratios. The magnitude of the chordwise and spanwise pressure distributions due to jet interference was greatest for rocket jets close to the wing and decreased as the jet was displaced farther from the wing. The design procedure for the rockets used is given in the appendix.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-893 , L-1614
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  • 17
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    Free Flight Observation of a Separated Turbulent Flow Including Heat Transfer Up to Mach 8.5 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A turbulent boundary layer separated by a forward-facing step was observed on the cylindrical portion of a hemisphere-cone-cylinder test vehicle. Tip blunting, producing a shear flow, was found to induce higher pressures on the cylindrical portion than were predicted from ballistic tunnel data of unblunted projectiles. An approximate method for predicting this blunt-body pressure distribution was hypothesized. These findings, along with the hypothesis, were substantiated by a wind tunnel test of a similar body. The peak pressure ratios of the separation were smaller in magnitude than flat plate theory predicted because of the effect of the shear flow. The decrement in heating of the separated flow, relative to the corresponding attached flow, was found to compare well with the expected results.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-278 , E-545
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  • 18
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    Heat Transfer to Surfaces of Finite Catalytic Activity in Frozen Dissociated Hypersonic Flow (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The heat transfer due to catalytic recombination of a partially dissociated diatomic gas along the surfaces of two-dimensional and axisymmetric bodies with finite catalytic efficiencies is studied analytically. An integral method is employed resulting in simple yet relatively complete solutions for the particular configurations considered. A closed form solution is derived which enables one to calculate atom mass-fraction distribution, therefore catalytic heat transfer distribution, along the surface of a flat plate in frozen compressible flow with and without transpiration. Numerical calculations are made to determine the atom mass-fraction distribution along an axisymmetric conical body with spherical nose in frozen hypersonic compressible flow. A simple solution based on a local similarity concept is found to be in good agreement with these numerical calculations. The conditions are given for which the local similarity solution is expected to be satisfactory. The limitations on the practical application of the analysis to the flight of the blunt bodies in the atmosphere are discussed. The use of boundary-layer theory and the assumption of frozen flow restrict application of the analysis to altitudes between about 150,000 and 250,000 feet.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-350 , A-425
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  • 19
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    Some Effects of Leading-Edge Sweep on Boundary-Layer Transition at Supersonic Speeds (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The effects of crossflow and shock strength on transition of the laminar boundary layer behind a swept leading edge have been investigated analytically and with the aid of available experimental data. An approximate method of determining the crossflow Reynolds number on a leading edge of circular cross section at supersonic speeds is presented. The applicability of the critical crossflow criterion described by Owen and Randall for transition on swept wings in subsonic flow was examined for the case of supersonic flow over swept circular cylinders. A wide range of applicability of the subsonic critical values is indicated. The corresponding magnitude of crossflow velocity necessary to cause instability on the surface of a swept wing at supersonic speeds was also calculated and found to be small. The effects of shock strength on transition caused by Tollmien-Schlichting type of instability are discussed briefly. Changes in local Reynolds number, due to shock strength, were found analytically to have considerably more effect on transition caused by Tollmien-Schlichting instability than on transition caused by crossflow instability. Changes in the mechanism controlling transition from Tollmien-Schlichting instability to crossflow instability were found to be possible as a wing is swept back and to result in large reductions in the length of laminar flow.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1075 , A-461
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  • 20
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    Boundary-Layer Transition on a Group of Blunt Nose Shapes at a Mach Number of 2.20 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An investigation has been made to study boundary-layer transition on six axisymmetrical blunt bodies of revolution. Model shapes were selected with respect to the degree of favorable pressure gradient over the model surface. Tests were conducted at a Mach number of 2.20 and over a range of free-stream Reynolds number per foot of about 1.4 x 10(exp 6) to 6.5 x 10(exp 6). The tests were made at an angle of attack of 0 deg. with zero heat transfer. For the hemisphere, the flow remained essentially laminar over the model surface length for the entire pressure range of the tests. For a strong favorable pressure gradient followed by any weak favorable, neutral, or adverse gradient, the tendency was for transition to occur at or immediately behind the shoulder. A single strip of three-dimensional roughness in the region of strong favorable pressure gradient did not fix transition on the models at the roughness location except at the maximum test pressures, whereas a second roughness strip added in a region of neutral or adverse pressure gradient did fix transition. Experimental pressure coefficients agreed closely with modified Newtonian theory except in the shoulder region.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-932 , L-1383
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  • 21
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    Estimate of Shock Standoff Distance Ahead of a General Stagnation Point (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The shock standoff distance ahead of a general rounded stagnation point has been estimated under the assumption of a constant-density-shock layer. It is found that, with the exception of almost-two-dimensional bodies with very strong shock waves, the present theoretical calculations and the experimental data of Zakkay and Visich for toroids are well represented by the relation Delta-3D/R(s) = ((Delta-ax sym)/(R(s))/(2/(K+1))) where Delta is the shock standoff distance, R(s),x is the smaller principal shock radius, and K is the ratio of the smaller to the larger of the principal shock radii.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1050 , E-1278
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  • 22
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    Laminar Heat-Transfer and Pressure Measurements at a Mach Number of 6 on Sharp and Blunt 15 deg Half-Angle Cones at Angles of Attack Up to 90 deg (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Two circular conical configurations having 15 deg half-angles were tested in laminar boundary layer at a Mach number of 6 and angles of attack up to 90 deg. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted nose with a bluntness ratio of 0.1428 and a fineness ratio of 1.66. Pressure measurements and schlieren pictures of the flow showed that near-conical flow existed up to an angle of attack of approximately 60 deg. At angles of attack above 70 deg high-pressure areas were present near the base and the bow shock wave was considerably curved. Comparison of the results with simply applied theories showed that on the stagnation line pressures may be predicted by Newtonian theory, and heat transfer by local yawed-cylinder theory based on the yaw angle of the windward generator and the local radius of the cone. Base effects increased the heat transfer in a region extending forward approximately 15 to 30 percent of the windward generator. Circumferential pressure distributions were higher than the corresponding Newtonian distribution and a better prediction was obtained by modifying the theory to match the pressure at 90 deg from the windward generator to that on the surface of the cone at an angle of attack of 0 deg. Circumferential heat-transfer distributions were predicted satisfactorily up to about 60 deg from the stagnation line by using Lees' heat-flux distribution based on the Newtonian pressure. The effects of nose bluntness at large angles of attack were very small in the region beyond two nose radii from the point of tangency.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-962 , L-1624
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  • 23
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    A Nonlinear Theory for Predicting the Effects of Unsteady Laminar, Turbulent, or Transitional Boundary Layers on the Attenuation of Shock Waves in a Shock Tube with Experimental Comparison (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The linearized attenuation theory of NACA Technical Note 3375 is modified in the following manner: (a) an unsteady compressible local skin-friction coefficient is employed rather than the equivalent steady-flow incompressible coefficient; (b) a nonlinear approach is used to permit application of the theory to large attenuations; and (c) transition effects are considered. Curves are presented for predicting attenuation for a shock pressure ratio up to 20 and a range of shock-tube Reynolds numbers. Comparison of theory and experimental data for shock-wave strengths between 1.5 and 10 over a wide range of Reynolds numbers shows good agreement with the nonlinear theory evaluated for a transition Reynolds number of 2.5 X 10(exp 5).
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TR-R-85
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  • 24
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    Local Aerodynamic Heat Transfer and Boundary-Layer Transition on Roughened Sphere-Ellipsoid Bodies at Mach Number 3.0 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A wind-tunnel investigation was made to determine heat-transfer distributions on three steel sphere-ellipsoid bodies with surface roughnesses of 5, 100, and 200 microinches. Tests were conducted in the Langley 9- by 6-foot thermal structures tunnel at a Mach number of 3.0, free-stream Reynolds numbers (based on model spherical diameter) of 4.25 x 10(exp 6) and 2.76 x l0(exp 6), and at a stagnation temperature of 650 F. Pressure distributions were obtained also on a fourth model. The results indicated that the combination of surface roughness and boundary-layer cooling tended to promote early transition and nullify the advantages attributable to the blunt shape of the model for reducing local temperatures. Good correlation between experimental heating rates and those calculated from laminar theory was achieved up to the start of boundary-layer transition. The correlation also was good with the values predicted by turbulent theory for surface stations downstream from the 45 deg. station.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-907 , L-1393
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  • 25
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    Configuration Factors for Exchange of Radiant Energy Between Axisymmetrical Sections of Cylinders, Cones, and Hemispheres and Their Bases (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Radiation-interchange configuration factors are derived for axisymmetrical sections of cylinders, cones, and hemispheres radiating internally to annular and circular sections of their bases and to other axisymmetrical sections. The general procedure of obtaining configuration factors is outlined and the results are presented in the form of equations, tables, and figures.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-944 , L-992
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  • 26
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    Analysis of Heat Transfer and Pressure Drop for a Gas Flowing Through a set of Multiple Parallel Flat Plates at High Temperatures (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Equations were derived representing heat transfer and pressure drop for a gas flowing in the passages of a heater composed of a series of parallel flat plates. The plates generated heat which was transferred to the flowing gas by convection. The relatively high temperature level of this system necessitated the consideration of heat transfer between the plates by radiation. The equations were solved on an IBM 704 computer, and results were obtained for hydrogen as the working fluid for a series of cases with a gas inlet temperature of 200 R, an exit temperature of 5000 0 R, and exit Mach numbers ranging from 0.2 to O.8. The length of the heater composed of the plates ranged from 2 to 4 feet, and the spacing between the plates was varied from 0.003 to 0.01 foot. Most of the results were for a five- plate heater, but results are also given for nine plates to show the effect of increasing the number of plates. The heat generation was assumed to be identical for each plate but was varied along the length of the plates. The axial variation of power used to obtain the results presented is the so-called "2/3-cosine variation." The boundaries surrounding the set of plates, and parallel to it, were assumed adiabatic, so that all the power generated in the plates went into heating the gas. The results are presented in plots of maximum plate and maximum adiabatic wall temperatures as functions of parameters proportional to f(L/D), for the case of both laminar and turbulent flow. Here f is the Fanning friction factor and (L/D) is the length to equivalent diameter ratio of the passages in the heater. The pressure drop through the heater is presented as a function of these same parameters, the exit Mach number, and the pressure at the exit of the heater.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1165 , E-1381
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  • 27
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    Experimental Study of Combined Forced and Free Laminar Convection in a Vertical Tube (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An apparatus was built to verify an analysis of combined forced and free convection in a vertical tube with uniform wall heat flux and to determine the limits of the analysis. The test section was electrically heated by resistance heating of the tube wall and was instrumented with thermocouples in such a way that detailed thermal entrance heat-transfer coefficients could be obtained for both upflow and downflow and any asymmetry in wall temperature could be detected. The experiments showed that fully developed heat-transfer results, predicted by a previous analysis, were confirmed over the range of Rayleigh numbers investigated. The concept of "locally fully developed" heat transfer was established. This concept involves the assumption that the fully developed heat-transfer analysis can be applied locally even though the Rayleigh number is varying along the tube because of physical-property variations with temperature. Thermal entrance region data were obtained for pure forced convection and for combined forced and free convection. The analysis of laminar pure forced convection in the thermal entrance region conducted by Siegel, Sparrow, and Hallman was experimentally confirmed. A transition to an eddy motion, indicated by a fluctuation in wall temperature was found in many of the upflow runs. A stability correlation was found. The fully developed Nusselt numbers in downflow were below those for pure forced convection but fell about 10 percent above the analytical curve. Quite large circumferential variations in wall temperature were observed in downflow as compaired with those encountered in upflow, and the fully developed Nussalt numbers reported are based on average wall temperatures determined by averaging the readings of two diametrically opposite wall thermocouples at each axial position. With larger heating rates in downflow the wall temperature distributions strongly suggested a cell flow near the bottom. At still larger heating rates the wall temperatures varied in a periodic way.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1104
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  • 28
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    An Investigation of the Natural Frequencies and Mode Shapes of Liquids in Oblate Spheroidal Tanks (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An experimental investigation was conducted to gain some understanding of the character of the free vibration modes of liquids In oblate spheroidal tanks applicable in missile and space vehicle systems, Measured natural frequencies were obtained for the lowest three or four, antisymmetric modes of oscillation as a function of the liquid depth for three orientations of each of several such tanks of different size and oblateness. The orientations considered were such that: (a) the equator of the spheroid was horizontal and oscillations were along a diameter of the circular liquid surface; (b) the equator of the spheroid was vertical and the oscillations were along the minor axis of the elliptical liquid surface; and (c) the equator of the spheroid was vertical and the oscillations were along the major axis of the elliptical liquid surface; The frequency data are presented as dimensionless parameters developed for each orientation to permit the application of the experimental results to the prediction of the natural frequencies of tanks of different size and oblateness. Photographs we re made of representative surface wave or mode, shapes for each orientation.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-904 , L-1264
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  • 29
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    Free-Flight Investigation of Heat Transfer to an Unswept Cylinder Subjected to an Incident Shock and Flow Interference from an Upstream Body at Mach Numbers up to 5.50 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Heat-transfer rates have been measured in free flight along the stagnation line of an unswept cylinder mounted transversely on an axial cylinder so that the shock wave from the hemispherical nose of the axial cylinder intersected the bow shock of the unswept transverse cylinder. Data were obtained at Mach numbers from 2.53 to 5.50 and at Reynolds numbers based on the transverse cylinder diameter from 1.00 x 10(exp 6) to 1.87 x 10(exp 6). Shadowgraph pictures made in a wind tunnel showed that the flow field was influenced by boundary-layer separation on the axial cylinder and by end effects on the transverse cylinder as well as by the intersecting shocks. Under these conditions, the measured heat-transfer rates had inconsistent variations both in magnitude and distribution which precluded separating the effects of these disturbances. The general magnitude of the measured heating rates at Mach numbers up to 3 was from 0.1 to 0.5 of the theoretical laminar heating rates along the stagnation line for an infinite unswept cylinder in undisturbed flow. At Mach numbers above 4 the measured heating rates were from 1.5 to 2 times the theoretical rates.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-988 , L-879
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  • 30
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    Hypersonic Viscous Shock Layer of Nonequilibrium Dissociating Gas (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The nonequilibrium chemical reaction of dissociation and recombination is studied theoretically for air in the viscous shock layer at the stagnation region af axisymmetric bodies. The flight regime considered is for speeds near satellite speed and for altitudes between 200,000 and 300,000 feet. The convective heat transfer to noncatalytic walls is obtained. The effects of nose radius, wall temperature, and flight altitude on the chemical state of the shock layer are studied. An analysis is also made on the simultaneous effect of nonequilibrium chemical reaction and air rarefaction on the shock layer thickness.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TR-R-109
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  • 31
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    Applicability of Mixing Length Theory to a Turbulent Vortex System (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The ability of mixing length theory to correlate vortex data is evaluated. Expressions are derived for eddy diffusivity by applying the techniques of von Karman and Prandtl which have been established for pipe flow. Total and static pressures were measured from the outer radius to the exhaust-nozzle radius of a vortex generator for a range of mass flows. These data are combined with Navier-Stokes solutions for this region of a compressible vortex to determine turbulent Reynolds numbers. The Reynolds number is related to Prandtl and Karman functions for various assumed boundary conditions, and the experimental data are used to determine the usefulness of these expressions. The following conclusions were reached: (1) Mixing length functions developed by applying von Karman's similarity hypothesis to vortex motion correlate the data better than do Prandtl functions obtained with the assumption that mixing length is proportional to radius. (2) Some of the expressions developed do not adequately represent the experimental data. (3) The data are correlated with acceptable scatter by evaluating the fluid radial inertia at the outer boundary and the shear stress at the inner boundary. The universal constant K was found to be 0.04 to 0.08, rather than the value of 0.4 which is accepted for rectilinear flow. (4) The data are best correlated by a modified Karman expression which includes an effect of radial inertia, as well as shear stress, on eddy diffusivity.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1051 , E-1145
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  • 32
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    Forces and Moments on Sphere-Cone Bodies in Newtonian Flow (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The static longitudinal aerodynamic characteristics of a family of sphere-cone combinations (fineness ratios from 1.0 to 6.0) were computed by means of Newtonian impact theory. The effects of angle of attack, fineness ratio, and center-of-gravity location are shown. The results indicate that, with the center of gravity at or near the center of volume, the sphere-cone combinations are statically stable at trim points that provide low to moderate lift-drag ratios. In general, the lift-drag ratio increased with increasing fineness ratio. As an example, with the center of gravity at the center of volume, the lift-drag ratio at trim was increased from approximately 0.05 to 0.56 by increasing the fineness ratio from 1.2 to 6.0.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1203
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  • 33
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    Transition of the Laminar Boundary Layer on a Delta Wing with 74 degree Sweep in Free Flight at Mach Numbers from 2.8 to 5.3 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The tests were conducted at Mach numbers from 2.8 to 5.3, with model surface temperatures small compared to boundary-layer recovery temperature. The effects of Mach number, temperature ratio, unit Reynolds number, leading-edge diameter, and angle of attack were investigated in an exploratory fashion. The effect of heat-transfer condition (i.e., wall temperature to total temperature ratio) and Mach number can not be separated explicitly in free-flight tests. However, the data of the present report, as well as those of NACA TN 3473, were found to be more consistent when plotted versus temperature ratio. Decreasing temperature ratio increased the transition Reynolds number. The effect of unit Reynolds number was small as was the effect of leading-edge diameter within the range tested. At small values of angle of attack, transition moved forward on the windward surface and rearward on the leeward surface. This trend was reversed at high angles of attack (6 deg to 18 deg). Possible reasons for this are the reduction of crossflow on the windward side and the influence of the lifting vortices on the leeward surface. When the transition results on the 740 delta wing were compared to data at similar test conditions for an unswept leading edge, the results bore out the results of earlier research at nearly zero heat transfer; namely, sweep causes a large reduction in the transition Reynolds number.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1066 , A-589
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  • 34
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    Interaction of Highly Underexpanded Jets with Simulated Lunar Surfaces (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Pressure distributions and erosion patterns on simulated lunar surfaces (hard and soft) and interference effects between the surface and two representative lunar vehicles (cylindrical and spherical) were obtained with cold-air jets at various descent heights and nozzle total-pressure ratios up to 288,000. Surface pressure distributions were dependent on both nozzle area ratio and, nozzle contour. Peak pressures obtained with a sonic nozzle agreed closely with those predicted theoretically for a near-sonic jet expanding into a vacuum. Short bell-shaped nozzles gave annular pressure distributions; the low center pressure resulted from the coalescence of shocks that originated within the nozzle. The high surface pressures were contained within a circle whose diameter was about 16 throat diameters, regardless of nozzle area ratio or contour. The peak pressure increased rapidly as the vehicle approached the surface; for example, at a descent height of 40 throat diameters the peak pressure was 0.4 percent of the chamber pressure, but increased to 6 percent at 13 throat diameters. The exhaust jet eroded a circular concave hole in white sand at descent heights from about 200 to 600 throat diameters. The hole diameter was about 225 throat diameters, while the depth was approximately 60 throat diameters. The sand particles, which formed a conical sheet at a semivertex angle of 50 deg, appeared to follow a ballistic trajectory and at no time struck the vehicle. An increase in pressure was measured on the base of the cylindrical lunar vehicle when it approached to within 14 throat diameters of the hard, flat surface. No interference effects were noted between the spherical model and the surface to descent heights as low as 8 throat diameters.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1095 , E-1300
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  • 35
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    Measurement of Flow Angularity at Supersonic and Hypersonic Speeds with the Use of a Conical Probe (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The characteristics of a 40 deg half-angle cone for measuring flow angularity were determined experimentally. Tests were conducted at a Mach number of 21 in helium with check points at Mach number 3.55 in air for angles of pitch up to 5 deg. The air tests confirmed theoretical indications of small or negligible Mach number and test-medium effects for the case of air and helium. The instrument is capable of measuring flow angularity at high Reynolds numbers and speeds greater than that necessary for shock attachment to within (+/-)1/3 deg.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-959 , L-1216
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  • 36
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    Heat Transfer and Boundary-Layer Transition on a Highly Polished Hemisphere-Cone in Free Flight at Mach Numbers Up to 3.14 and Reynolds Numbers Up to 24 x 10(exp 6) (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A highly polished hemisphere-cone having a ratio of nose radius to base radius of 0.74 and a half-angle of 14.5 was flight tested at Mach numbers up to 4.70. Temperature and pressure data were obtained at Mach numbers up to 3.14 and a free-stream Reynolds number of 24 x 10(exp 6) based on body diameter. The nose of the model had a surface roughness of 2 to 5 microinches as measured with an interferometer. The measured Stanton numbers were in good agreement with theory. Transition Reynolds numbers based on the laminar boundary-layer momentum thickness at transition ranged from 2,190 to 794. Comparison with results from previous tests of blunt shapes having a surface roughness of 20 to 40 microinches showed that the high degree of polish was instrumental in delaying the transition from laminar to turbulent flow.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-955 , L-1701
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  • 37
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    Radiation Emission Effects of the Equilibrium Boundary Layer in the Stagnation Region (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The laminar compressible boundary layer in chemical equilibrium is analysed to show the effect of emission of radiation by the boundary layer on solutions of the energy equation and on the resulting heat transfer. Solutions are obtained at one flight condition, but for several nose radii, in a regime where absorption is negligible. The consequent effects on heat transfer and boundary-layer thickness are determined. The concept of a separate boundary layer and shock layer is discussed in the light of the results obtained.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1031 , A-509
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  • 38
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    Analysis of Multipoint-Multitime Correlations and Diffusion in Decaying Homogeneous Turbulence (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Two-point, two-time correlation equations are obtained by considering the Navier-Stokes equations for two points in a fluid at two time. By neglecting the triple correlations in the equations, a solution is obtained for the final period of decay. The analysis is extended to earlier times by considering three points at three different times. The set of equations is made determinate by neglecting the quadruple correlations in comparison with the triple correlations. The diffusion of particles from a source in a decaying turbulent field is calculated approximately by assuming that the velocity fluctuations are small.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TR-R-96
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  • 39
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    Calculation of Flow Fields from Bow-Wave Profiles for the Downstream Region of Blunt-Nosed Circular Cylinders in Axial Hypersonic Flight (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A method by which known bow-wave profiles may be analyzed to give the flow fields around blunt-nosed cylinders in axial hypersonic flow is presented. In the method, the assumption is made that the pressure distribution curve in a transverse plane is similar to that given by blast- wave theory. Numerical analysis based on the one-dimensional energy and continuity equations then leads to distributions of all the flow variables in the cross section, for either a perfect gas or a real gas. The entire flow field need not be solved. Attention can be confined to any desired station. The critical question is the validity of the above assumption. It is tested for the case of a hemisphere cylinder in flight at 20,000 ft/sec. The flow is analyzed for three stations along the cylindrical afterbody, and found to compare very closely with the results of an exact (inviscid) solution. The assumed form of the pressure distribution occurs at stations as close as 1.2 diameters to the body nose. However, it is suggested that the assumption may not apply this far forward in general, particularly when bodies of nonsmooth contour are considered.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1147 , A-493
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  • 40
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    Heat Transfer at the Reattachment Zone of Separated Laminar Boundary Layers (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The flow and heat transfer are analyzed at the reattachment zone of two-dimensional separated laminar boundary layers. The fluid is considered to be flowing normal to the wall at reattachment. An approximate expression is derived for the heat transfer in the reattachment region and a calculated value is compared with an experimental measurement.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1072 , A-561
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  • 41
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    The Effect of Air Bleed on the Heat Transfer and Pressure Distribution on 30 Conical Flares at a Mach Number of 6.8 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The effect of a boundary-layer bleed at the start of a 30 deg half-angle flare upon the shape of the flow boundaries, the pressure distribution on the flare, and the heat transfer to the flare was studied at a Mach number of 6.8. The forebody was an ogive cylinder. Test Reynolds numbers, based on forebody length, ranged from 1 x 10(exp 6) to 7.4 x 10(exp 6). Schlieren photographs showed the effect produced upon the flow boundaries by varying the dimensions of the bleed in both the radial and axial directions and by blunting the lip at the leading edge of the flare. The heat transfer and pressure distribution on the flare were correlated with the shape and nature of the flow boundaries.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TM-X-439
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  • 42
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    Base Flow Characteristics for Several Four-Clustered Rocket Configurations at Mach Numbers from 2.0 to 3.5 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A generalized study of base flow phenomena has been conducted with four 500-pound-thrust JP-4 fuel-liquid-oxygen rocket motors installed in the base of a 12-inch-diameter cylindrical model. Data were obtained over a Mach number and nozzle pressure ratio range of 2.0 to 3.5 and 340 to 600, respectively. Base heat flux, gas temperature, and pressure were highest in the center of the cluster core and decreased in a radial direction. Although a maximum heat flux of 93 Btu per square foot per second was measured within the cluster core, peripheral heat fluxes were low, averaging about 5 Btu per square foot per second for all configurations. Generally base heat flux was found to be independent of Mach number over the range investigated. Base heat flux within the cluster core was decreased by increasing motor spacing, motor extension, a combination of increasing nozzle area ratio and decreasing exit angle and gimbaling the two side engines. Small amounts of nitrogen injected within the cluster core sharply reduced core heat flux.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1093 , E-1241
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